1. Field of the Invention
This invention relates generally to a gas turbine engine shroud, and particularly relates to a uniformly cooled and pressure balanced segmented shroud wherein each shroud segment continuously spans both the high pressure turbine blades and the low pressure turbine blades. This design eliminates a row of stationary vanes between the rotating blades thereby providing a large reduction in weight, significant cost savings and increased performance through reduced cooling air requirements.
2. Description of Prior Developments
The primary function of a gas turbine engine shroud is to provide a contoured annular surface along the exhaust gas outer flowpath and to define as small a clearance as possible with the tips of the rotating turbine blades. Maintaining this small clearance is necessary to minimize the escape of exhaust gas between the blade tips and the outer flowpath surface. The radial clearance between the rotating blade tips and the stationary shroud has a significant effect on turbine efficiency, with small clearance providing greater efficiency.
The effect of blade tip clearance on turbine efficiency and performance is most significant on the high reaction gas turbine applications in which the present invention is used. The tighter the clearance gap can be maintained, the better the performance of the turbine. Therefore, much effort is placed in the design of the shroud as well as its shroud support to provide maximum control over the radial position of the shroud, as the radial position of the shroud defines the blade tip clearance.
Since the minimum clearance between the shroud and the blades, i.e. the pinch-point, normally occurs during transient operation, it is of critical importance to control the transient response of the shroud support in order to maintain acceptable blade tip clearance levels at steady state operating conditions. Ideally, the stator response should match the rotor transient response in order to achieve minimum steady-state clearances and improve engine performance.
To achieve good engine performance, it is also necessary to maintain the shroud and its shroud support as round as possible. Non-uniform mechanical and/or thermal radial loads which tend to distort the shroud support and the shroud may cause local rubbing on the shroud by the blade tips. This creates non-uniform shroud wear and associated blade tip loss and results in degraded engine performance.
The shroud support design shown in FIG. 1 is typical of known conventional designs. The clearance control or support rings 10, 12 formed on the engine case 14 are heated and cooled by cooling air circuits which direct the cooling air tangentially within channels formed between the clearance control rings. The high pressure turbine shroud 18 is separate and axially spaced from the low pressure turbine shroud 20. The free ends of the high pressure turbine blades 22 and the low pressure turbine blades 24 define clearance gaps 25 with the respective shrouds 18, 20.
Testing of this conventional design has revealed circumferential temperature gradients exceeding 80.degree. F. This temperature variation is believed to be primarily due to the under cowl environment and leakage of cooling air around various pipe fittings 16. Such temperature gradients may drive open the blade tip clearance gaps 25 by 0.008 inch after blade tip rubbing. This is a significant penalty since steady state clearances are generally in the range of 0.015-0.020 inch.
A major concern in the design of any shroud system is its ability to use cooling air effectively and to reduce parasitic leakage of this air. Current high pressure turbine designs are cooled using compressor discharge air routed around the combustor and nozzle outer support bands. Leakage of this air to the exhaust gas flowpath is typically controlled by using thin sheet metal shim seals between shroud segment ends. Such conventional shroud designs allow full shroud coolant pressure to leak across these seals. This leakage is represented in FIG. 1 by directional arrows 23.
More recent designs, such as that shown in FIG. 2, have incorporated continuous 360.degree. impingement baffles 26, thereby reducing the pressure differential across the shroud end seals 21. This lower pressure differential results in reduced coolant leakage. The 360.degree. impingement baffle design, however, is not adaptable to a segmented shroud hanger configuration such as that schematically depicted in FIG. 2(a). This can be a drawback as it is desirable to form the shroud hangers 19 as a series of circumferentially spaced segments which prevent the non-uniformly heated flowpath shrouds 18 from influencing the temperature of the shroud support which is preferably formed as a continuous 360.degree. support ring 12. In this manner, the segmented shroud hanger thermally isolates the shroud from the support ring 12.
Accordingly, a need exists for a segmented gas turbine engine shroud which maintains a close, circumferentially uniform clearance with respect to the rotating turbine blades during both transient and steady state engine operating conditions.
A further need exists for a gas turbine engine shroud support which is evenly circumferentially heated and cooled so that circumferential temperature gradients are avoided and so that the attached shrouds are maintained as close to round as possible at all times.
Yet another need exists for a gas turbine engine shroud which effectively uses cooling air by reducing pressure differentials across the shroud seals thereby reducing parasitic leakage of the cooling air.
Another object of the invention is to control and uniformly maintain the heat transfer coefficients along the shroud support, and particularly along the annular radial flanges which form the three shroud support position control rings.
Another object of the invention is to control the pressure adjacent and between the shroud support and the segmented shroud so that radial loads on these members are minimized or eliminated.
Another object of the invention is to provide a shroud which spans two adjacent rotors and provides blade tip clearance control to both. Use of separate shrouds for each rotor would result in more component parts, joints and greater leakage of cooling air through the joints.
Still another object of the invention is to facilitate the assembly and disassembly of a segmented gas turbine engine shroud to and from its hangers and shroud support member.